Automatic approach landing and go-around pitch axis control system for aircraft

ABSTRACT

The invention relates to a system utilizing an exponential control law for glide slope capture and flare. The capture maneuver from above or below the beam, is a function of decreasing glide slope beam error in the same manner that the flare maneuver is a function of decreasing altitude above the runway. The present autopilot approach coupler is an altitude rate command system which provides switchless signal processing during glide slope capture, tracking and flare, and in addition, provides automatic go-around control from any altitude during the approach.

United States Patent Simpson et a1. Apr. 2, 1974 [54] AUTOMATIC APPROACHLANDING AND 3,291,421 12/1966 Kramer et al 244/77 A GO AROUND PITCH AXISCONTROL 3,601,339 8/1971 Watson 244/77 A 3,327,973 6/1967 Kramer et al244/77 A SYSTEM FOR AIRCRAFT 3,658,280 4/1972 McDonnell 244/77 D [75]Inventors: Robert D. Simpson; Jimmie H. 3,578,269 5/1971 Kramer ct81,..." 244 77 A Boone, both of Bellevue; Gary A. 3,447,765 6/1969Doniger et a1. 244/77 A Chenkovich, Seattle, all of Wash. PrimaryExaminer-Duane A. Reger [73] Assgnee' g Eating Company Seattle AssistantExaminerStephen G. Kunin as Attorney, Agent, or Firm-Conrad O, Gardner;Glenn [22] Filed: Jan. 31, 1972 Orlob [2111, Appl. N0.: 221,958

[57] ABSTRACT [52] US Cl 244/77 A 235/5022 318/583 The invention relatesto a system utilizing an exponen- 318/584 543/108 tial control law forglide slope capture and flare. The 511 lm. c1. B64c 13/18 capturemaneuver above or below the beam is a [58] Field 0Search 235/150 22.244/77 A 77 D function of decreasing glide slope beam error in the244/77 318/583 343/5 7 same manner that the flare maneuver is a functionof 108 decreasing altitude above the runway. The present autopilotapproach coupler is an altitude rate command [56] References Citedsystem which provides switchless signal processing during glide slopecapture, tracking and flare, and in UNITED STATES PATENTS addition,provides automatic go-around control from Cvzllztssegnet al 12 anyaltitude during the approach 3,652,835 3/1972 Devlin et al. 244/77 A X11 Claims, 13 Drawing Figures ,4z 7/ 7Z/DE FA 7E 225720702 M P/TCH F475DETECTOE E EOE DE TEE TOE 4! 7/ 71/05 ,4501 5 TEFFAV/V 25 TEL 702.i'l/AFACE FEEDBACK CONTROL ill/( 54855 ELEM FDBK ELEV. EFEOE h COMMANDREFERENCE Tia/6A TOE Pi IPES'PONSE SHEEI 5 [1F 9 h COMM/1ND A29REFERENCE n,

N P1564 T/ONA L 4MP! lF/EE PATENFEU 2 SHEET 7 [IF 9 PAIENIEDAPR 2 I974SHEET 8 BF 9 Nommm EE AUTOMATIC APPROACH LANDING AND (IO-AROUND llllTCHAXIS CONTROL SYSTEM FOR AIRCRAFT This invention relates to signalprocessing for automatic approach and landing of an aircraft and moreparticularly relates to an improved system for automatically controllingthe pitch axis of an aircraft during an automatic approach and landingsequence.

Prior art control systems which fly an aircraft close to the landingrunway and thereafter approach that runway and flare-out for touchdownare available as exemplified by U. 5. Pat. No. 3,327,973 to Kramer etal. However, such a system utilizes a landing control law which flightthe flight crew to preselect a reference light path approach by theaircraft to the runway which reference flight path may or may not beoptimized for the desired flight path. During normal flight and prior toapproach for automatic landing, the system in the above referencedKramer et a]. patent utilizes an automatic system for controlling theelevators and thus the pitch attitude of the aircraft. Engagement of theautomatic landing system with preselected reference flight path in theaforementioned manner by the flight crew which results in less thanoptimum flight path acquisition further results in abrupt movements ofthe aircraft and large initial flight path errors. Such abrupt movementsare highly objectionable in commercial aviation since causing passengeralarm and discomfort, and also very importantly, large initial flightpath errors limit the ability of system performance at low altituderesulting in consequent deterioration of noise abatement procedures.

It is accordingly an object of the present invention to provide a pitchaxis control system for providing smooth and accurate acquisition of theglide slope beam to prevent large errors in flight path at extremely lowaltitudes.

It is a further object of this invention to provide a pitch axis controlsystem permitting capture of glide slope error independent of glideslope angle and approach speed for various altitudes of glide slopecapture.

It is yet another object of this invention to provide a pitch axiscontrol system having automatic nose lowering characteristics aftertouchdown not requiring system disengagement after touchdown.

It is still a further object of the present invention to eliminate useof programmed test unit (ATU) in autopilot glide slope flare couplerdesign for alternately switching in and out signals, sensors, andsubmodes to assure system integrity in favor of performing a totalintegrity test through the use of accelerations, errors, and dynamicoperation of the system during an initial maneuver such as the glideslope capture maneuver.

It is yet a further object of the present invention to provide in anautopilot control system, a synchronized automatic exponential captureof the glide slope error independent of glide slope angle.

The above and further objects are achieved in the present invention bysignal processing means for coupling control signals to utilizationmeans, e.g., the pitch axis control system which processing meansemploys a single set of control laws for signal processing during glideslope capture, tracking and flare maneuvers.

Other objects, features and advantages of the present invention willbecome apparent from the following description read on the accompanyingdrawings, wherein:

FIG. I corresponds to FIG. I of U. S. Pat. No. 3,327,973 which isincluded for ready reference to assist in comparison of the system ofthe present invention with this prior art;

FIG. 2 is a diagram useful in showing geometric rela tionships of theaircraft in relation to the ground for developing the equations offlight path control of the system of the present invention;

FIG. 3 is a block diagram showing signal processing utilized to deriveaircraft path command signals in accordance with the system of thepresent invention deemed helpful in further development of the equationsof flight path control of the system;

FIG. 4 is a graph showing commanded flight path as a function of a timesubsequent to glide slope capture further helpful in understanding thederived equation for commanded flight path;

FIG. 5 is a block diagram representative of signal processing foraircraft short period damping satisfied in the system of the presentinvention;

FIG. 6 is a block diagram illustrative of signal processing of thepresent system during the flare command phase of landing of theaircraft;

FIG. 7 is a block diagram showing system signal processing during thego-around phase;

FIG. 8 is a block diagram similar to FIG. 7 however in a more detailedaircraft environment;

FIG. 9 is an exemplary circuit embodiment of the goaround systems ofFIGS. 7 and 8;

FIG. 10 is a block diagrammatic representation of an embodiment of apitch axis control system according to the present invention; and

FIG. 11 shows in more detail the acceleration normal to flight pathdetector of FIG. 8;

FIG. 12 is an actual plot showing actual flight path compared to desiredflight path illustrative of the present pitch axis control systemperformance during glide slope capture tracking and go-around; and

FIG. 13 is an actual plot showing pitch axis control system performanceincluding flare, touchdown and automatic nose lowering.

Turning now to the system of FIG. 1 which is representative of the priorart, a comparison therewith will be made with the system of FIG. 8 whichis illustrative of the system of the present invention to bring out thefeatures of the present system. The features of the present system maythen become focussed upon and appreciated in the subsequent analysisfrom a signal processing standpoint and later system embodimentdescription which further explain and amplify how these features andresultant advantages are achieved in accordance with the present pitchaxis control system.

The system of FIG. 1 provides a synchronized automatic capture of theglide slope error which depends upon pilot initiated computer inputinformation based upon glide slope angle, approach speed, etc., which isonly optimized for one set of environmental or airplane conditions andfor one capture altitude while the present system of FIG. 10 provides asynchronized automatic exponential capture of the glide slope errorindependent of glide slope angle, approach speed, wind conditions, etc.,which allows optimum performance under substantially any environmentalor aircraft conditions for substantially any altitude of glide slopecapture.

The system of FIG. 1 utilizes pitch attitude for minor loop stabilityresulting in looser control of the desired flight path in the presenceof wind and tends towards less reliability due to the added sensor whilethe present system of FIG. 10 does not require the use of a pitchattitude source for minor loop stability thus allowing for more accuratecontrol along the desired flight path and also eliminating the extrasensor reliability.

The system of FIG. 1 can be seen to utilize a vertical velocity computerto derive velocity errors relative to true vertical, not the desiredflight path. The system also requires a longitudinal accelerometer foroptimum compensation for wind conditions or airplane speed bleeds. Thepresent system of FIG. 10 in contrast uses a normal accelerometerhowever tilted relative to the aircraft body axis (see FIG. 11 for moredetail) to provide instantaneous normal velocity errors relative to thedesired flight path and to compensate for anylo n gi tremely lowaltitudes, and allows automatic goarounds even after the aircrafttouches down.

d. The present system circuit design is such that no i u of e -a unemelenqsgnree lt a increased sink rate of the airplane after initiationof go-around.

The system of FIG. 1 does not provide automatic nose lowering aftertouchdown. The pilot is required to disengage the autopilot aftertouchdown and lower the nose to the ground manually prior to braking theaircraft while the present system of FIG. provides automatic noselowering after touchdown which allows the pilot to leave the systemengaged after touchdown and puts in nose down elevator to hold theaircraft on the ground after touchdown.

The present pitch axis system circuit embodiment implements thefollowing four control equations:

, I 29f mm ipet .fsnsriens.

tudinal acceleration errors due to environmental conditions or aircraftspeed bleeds. The present system further eliminates the need for alongitudinal accelerometer to compensate for these errors.

The system of FIG. 1 utilizes a vertical velocity computer whichcomputes the vertical velocity and does not contain pitch rateinformation requiring both pitch rate and pitch attitude for minor loopstability, and further requires the monitoring of these sensors forautomatic landings. The present system uses a normal accelerometermounted forward of the aircraft center of gravity to provide a signalproportional to pitch acceleration which is passed through a lag filterto provide a pitch rate signal. The present system thus eliminates thepitch rate gyro as a critical sensor thus facilitating easier monitoringof the system.

The system of FIG. 1 utilizes a fixed vertical beam sensor switch pointdetector which is optimized for only one glide slope capture altitudeand is much less acceptable for lower altitude glide slope captureswhile the present system of FIG. 10 utilizes a vertical beam sensor(switch point detector) that is downstream of the glide slope gainprogrammer. This allows optimum glide slope captures at substantiallyany altitude by varying the glide slope capture point inversely withaltitude which is advantageous in noise abatement type approaches.

The present system embodiment of FIG. 10 provides an automatic go-aroundcommand as does the system of FIG. 1, however the system of FIG. 10utilizes the same circuitry already utilized in FIG. 10 to perform otherILS coupling functions and has the following features and functionaladvantages over the system of FIG. 1:

a. If go-around circuitry fails, the system of FIG. 10

will flare the airplane allowing time for pilot correction at extremelylow altitudes.

b. In the present system, the initial go-around command is independentof the final go-around command thereby allowing the aircraft to initiategoaround and assume positive rate of climb even if final command hasfailed.

0. The flare command is not inhibited by go-around which additionallyreduces altitude loss at exg ffhe flare commandf unction l2 [h+hbias-l-Kdie .iJbEJtQitQyttifunszti a 955mm.

FIG. 2 showing aircraft relative position, viz, geometric representationof glide slope geometry and FIG. 3 showing in block form aircraft pathcommand signal processing may now be considered in developing the patherror and then path command signal terms of the present pitch axiscontrol system where 5 angular error of airplane from glide slope centeras sensed by glide slope error detector 0 glide slope center referenceangle h distance from glide slope center and airplane receiving antennaperpendicular to glide slope center h marries fro airplin ififiigafiinfia and ground perpendicular to glide slope center h distance fromglide slope center and ground perpendicular to glide slope center hvertical distance from glide slope center and amend d h= (hm) hl 1 1+h mX? if h r hd-h o YI A-5;-e2/J7.3 for (h +h)h 57.3h

i X+Aw X+Ax) but 'k tan 9 Ax=h tan 6 But since Kvz(t) =h(t) dh(t) o Nowmapping from time domain to S-plane and replacing? h +h 1 with I o forsimplicity where h K K h path ermr=% (sw -j s+ j)- SH) For zero patherror, the path commanded relative to u bstit uting into h(z) the glideslope zero plane is defined by: (aw [z+ 2 1/2 101 8+1 h(t)=I0 fl le-( 11n Sz+ s l -,W in c t 7 K K The above path command equation h in termsof This is in {He time constants, natural frequency and damping is seenW W, V. c, ,1 r A, to result in a glide slope capture which is alwaysexpo- -H) nential (see FIG. 4) and which is always entered tan- (s) S+2zwnS+ n gentially determined by 1 Turning now to the portion of thesystem providing .where:

K short period damping shown in FIG. 5 the surface comg Z mand equationsare developed in the following: 1 1 1 and z=damping; wz: naturalfrequency Lmfm t (0 No+L9)K21-2+9o 3+Ls IOZO 'lieiic? I a n as 1 d- OzNo 2 2+ 2 2+ o 3+ So L c M s) [:S +2Q0) S+ n l Q n n Lemme V K T LeKgTg9K +L I Then li fs) is irithe form:

L =0 For Airplane Damping Satisfied:

NO, 9 and 6 at -bt] a are very small and can be neglected, therefore:

'imizmki 7 d For flare command, the signal processing elements of Fairplane to fly a programmed rate of climb. The prothe system are shownin FIG. 6 which results in flare grammed rate of climb is generated inthe form of first command signals derived as follows: H g and secondsignal components as follows:

I I l. The first signal component generated in the goerrur o+ commandSince around portion of the system as shown in FIG. 8 where commands aclimb rate of 300 feet per minute. The .K fi -t/18 h error signal whichis proportional to elevator I -F bm 515 for ia:)S command, can be seento be composed of three K fi -ms m r et t goemumi wh sk a trnnnrmrl 1=f[h+h ,,,+K he-* 10 Prior to Flare K hFLl/fl l 1. 6/8 displacement (shortterm 11 command) proerror i I e g l=fi+ii h= -11}, for zero path errorgramIhed to Zero at 65 feet K I'i i p K [Clio/S integral (dh commandreference) 6 1 7 response amping terms) K. 15 Dian; Flare K he K l.flare command (held at zero out ut until an altien-or bia s 1=Ah+j:[h+hbias+K6h tude of pp t y '5 feet) p 2. 6/8 integral (h commandreference) 3. h response (damping terms) 2 2. The second signalcomponent generated by the goaround portion of the system shown in FIG.7 is the term which actually causes the aircraft to perform And for zeropath error:

an automatic go-around. This term (h command Turning now to 7, thefollowing derivations reference) is proportional to the aircraft rate ofshow how go-around equations representative of these sink h the i ft iConducting an approach Signals are developed by the Portion of theSystem since when the aircraft is flying zero glide slope shown in FIG.7: a error (on glide slope centerline the output of the =li Ii h errorcommand O- h command ho While the system portion ho n i FIG, 7 and thintegrator circuit 9 must be equal and opposite to above equationsdefine the go-around command signals the h" response of the dampingterms of lag filter generated the explanation which follows inconnection 17 t null ut th 11" rr r and fly a Zer elevator with FIGS. 8and 9 will further serve to explain in a command. This h commandreference is a fly more physical sense and in a complete circuit sched nmmand S0 that the aircraft is descending matic respectively how thego-around function is a approximately 600 t 700 feet p r minute on the ahie d i h i f i m centerline of the glide slope. When the pilotinitiates an automatic go-around by pushing the goaround switch 24 (seeFIG. 8), two events occur: first, the glide slope displacement andintegral input paths are removed by switch 5,, so that no ref- From thepreceding block diagram and discussion, it should be noted that theautomatic go-around command used in the present autopilot approachsystem is not automatically initiated but requires pilot activation 6fthe go around Switch f FIG 8 (correspondingly erence to the glide slopecenterline is maintained i hi means 24 f FIG. 7 Comprising a Switch)during the go-around. This in itself does not cause which is preferablylocated on the throttle levers. If any gd'arouhd command to be generatedbut during an approach, the flight crew decides that condi- Causes the'y maintain an hold tions are not adequate to continue the approach,e.g., mand (afiXed output on the integrator Circuit Since traffic on therunway, or inadequate visibility for landthe input t0 the integrator isZero) Prior to flare 0r ing, the pilot can initiate an automaticgo-around by inif in flare 1655 than an altitude of about 53 f creasingthe thrust and activating the go-around switch. to ntinue to Hate theaircraft due to the flare This action will cause the autopilot tocommand the 7 command. The second event occurs simultaneously with theclosing of a resistive circuit path which washes out the glide slopeintegrator generated signal, It command reference. Since the output ofthe integrator circuit is a fly down command, washing out or eliminationof the integrator output signal is representative of a fly up commandhaving a time constant determined by the RC network formed by theswitched resistive circuit path and the capacitor providing theintegrator feedback. For a Boeing Airplane Company type 747 aircraft,this time constant equals approximately 4.5 seconds but is dependentupon the particular aircraft characteristics. This function causes theaircraft to break its rate of sink. In addition as can be seen in FIG.8, a voltage bias is summed in through a resistive network (not shown)to cause the aircraft to initially seek a climb rate of 300 feet perminute.

A second phase in go-around occurs after closing of switches S and Swhen the aircrafts flaps are raised to less than 23 to provide thego-around flap setting thereby switching in an additional gain path fromthe voltage bias and causing the lag circuit to command an additional700 feet per minute climb rate for a total command rate of 1,000 feetper minute.

An actual exemplary embodiment of the go-around system of FIGS. 7 and 8is shown in FIG. 9. If the goaround is initiated below 53 feet (flareregion) or just prior to 53 feet and the aircraft enters the flareregion, the flare computer will also command a decrease in rate of sinkwhich aids the go-around command and allows the automatic go-around tobe used safely at very low altitudes including after touchdown.Automatic nose lowering after touchdown is provided in the present pitchaxis control system which system does not uti lize pitch attitude as adamping term but in which primary damping is dependent on altitude rate.At touchdown, the flare command is requesting a sink rate of 2.5 to 3feet per second. When the aircraft lands, the

aircraft sink rate is reduced to zero in a very short time intervalwhich produces an error between the actual sink rate of zero and thecommanded sink rate of 2.5 to 3 feet per second. This results in a nosedown elevator command effort for providing an increase in sink rate to2.5 to 3 feet per second. The present pitch axis control system dampingpermits this maneuver in a controlled manner. Prior art systems whichutilize pitch attitude for damping cannot generate sufficient altituderate error at touchdown to lower the nose, hence the pilot mustdisconnect the control system and lower the nose manually.

Turning now to FIG. 10, there is shown the complete control system whichprovides the several functions, e.g., flare command, go-around, etc.,already separately discussed. In the following discussion referencenumerals corresponding to those used earlier will be used to identifycorresponding elements of the system.

In the system of FIG. 10, between the system output terminal and thesumming junction 10 there is coupled a negative feedback loop. Thisnegative feedback loop comprises the glide slope integrator 9 connectedthrough switching means 12 comprising a relay switch in the positionshown and series gain 140 for providing a synchronizing path. Thissynchronizing path provides two functions when operating in thesynchronizing 10 model The first function is for reducing signalspresent at the system output terminal 15 to reference potential (zero)by driving glide slope integrator circuit 9 so that the output signalvoltage of integrator circuit 9 is substantially equal and opposite tothe sum of the remaining signal voltages at summingjunction 10. In thismanner, the pitch axis control system output signal at computer systemoutput terminal 15 is maintained at reference potential (zero voltagelevel) to assure that no undesirable aircraft maneuver is experienced atthe time of engagement of the automatic approach and landing computer ofthe present pitch axis control system. The second function of thesynchronizing path is for providing glide slope capture initialconditions so that when the present automatic approach and landing pitchaxis control system is engaged by closing switching means 12 to thedotted line position, the present system will maneuver the aircraft ontothe glide slope zero plane. This function is accomplished in a uniqueand novel manner without having to switch in a separate signalgenerating means and by utilizing the same control laws previouslyderived which also provide the glide slope zero plane tracking. Sincethe glide slope integrator circuit 9 has stored at its output, a signalvoltage which is equal and opposite to the sum of all other signalvoltages appearing at the input of summing junction 10, and, for a glideslope capture from a point below the glide slope zero plane, this storedoutput contains a signal which is equal and opposite to the fly upcommand from the glide slope error detector 4 through the variable glideslope gain programmer circuit ll. Circuit 11 comprises means well-knownin the art for multiplying two variables together, e.g., pulse widthmodulated shunt switching means. At a fixed error signal level from theglide slope zero plane, the vertical beam sensor 66, threshold isexceeded causing switching means 12 to transfer and thus removing theoutput signal at terminal 15 from the input summing junction 8 of theglide slope integrator. The synchronizing path is removed by this actionand the glide slope integrator signal at this instant in time is fixedand can no longer change to drive the output signal at terminal 15 tozero for any change in the remaining input signals to summing junction10. As the aircraft continues to fly toward the glide slope zero plane(see FIG. 2), the fly up command from glide slope error detector 4 is reduced in magnitude thus creating an error signal at systern outputterminal 15 in a fly down command direction which comprises the storedfly down signal from glide slope integrator 9 and the decreased fly upsignal from glide slope error detector 4. The fly down command errorsignal at the output terminal 15 causes the elevator control system tocause displacement in a direction causing the aircraft to descend. Thisdisplacement of elevator surfaces in the control system is coupled bysurface feedback measuring means 16 to null the system output commandsignal voltage at system output terminal 15. The aircraft rate ofdescent signal voltage provided by altitude rate detector 2 and theaircraft rate of acceleration signal voltage provided by detector means1 (comprising an accelerometer having sensitive axis mounted normal tothe desired flight path) shown in FIG. 10 (and in complete detail inFIG. 11) sense that the aircraft is descending and these two signalvoltages are summed and coupled through lag filter means 17 (comprisinga low pass lag filter, e.g., a resistor in parallel with a capacitor infeedback circuit of an operational amplifier) to produce a signal whichis referenced to the aircraft flight path for short period maneuveringand to the aircraft vertical rate of descent for long term maneuvering.This uniquely derived signal is obtained by combining at adder 135(comprising, e.g., a summing junction): higher frequency signalcomponents from an accelerometer 1 which is tilted physically in theaircraft such that its sensing axis is disposed perpendicular to theflight path of the aircraft and which is positioned forward of thecenter of gravity of the aircraft (as shown in FIG. 11) which transmitsthese higher frequency signal components through high pass filtercircuit 131 and summing resistor 132, and lower frequency componentsfrom an altitude rate signal source 2 which is reference to verticalrate of descent. As the aircraft descends, the output signal from lagfilter 17 having the above higher and lower frequency signal componentsis representative of a fly down response in the system or a deviationfrom the aircraft flight path available through the circuit path coupledto junction to null or cancel the signal voltage representative ofcommanded deviation from the' aircraft flight path 119.

The results of the above described method of acquiring the glide slopezero plane is a fly down (or fly up if approaching from above the glideslope zero plane) altitude rate command signal voltage proportional tothe error between the stored glide slope error signal voltage at theoutput of glide slope integrator circuit 9 and the actual glide slopeerror signal voltage generated by glide slope error detector 4. In thismanner this unique feature of the present pitch axis control systemprovides a means for acquiring the glide slope zero plane which issubstantially independent of external factors such as aircraft speed,glide slope zero plane angles, and glide slope error signal gradients.The above feature is accomplished by utilizing only one signalgenerating means for both glide slope capture and tracking functions.

The above described pitch axis control system provides a flight pathcommand signal at the system output terminal which positions theaircraft on a flight path to exponentially acquire the glide slope zeroplane, and it will be further noted that the closing of the switch 12(to the position shown by the dotted line) also couples in seriescircuit path glide slope gain programmer circuit 11 and gain 141 betweenglide slope error detector circuit 4 and summing junction 8 therebyproviding a means for varying the stored glide slope error signalpresent at the output of glide slope integrator 9 during the glide slopeacquisition maneuver and subsequent glide slope zero plane tracking tothereby eliminate er rors developed in the flight path command signalpresent at system output terminal 15 and as a consequence cause theaircraft to acquire and track the zero plane of the glide slope errorsignal. The glide slope integrator output signal voltage from integratorcircuit 9 at this time is proportional to but of opposite polarity tothe descent rate signal voltage of the aircraft at low pass lag filter17 which relationship is required to maintain the glide slope errorsignal voltage at glide slope error detector 4 equal to zero.

The damping terms for the pitch axis control system of FIG. 10 arederived in a novel and unique manner by mounting of the accelerometer 1in the manner shown in FIG. 11, viz., normal to the flight path andforward of the aircraft center of rotation such that the output ofaccelerometer circuit 1 comprises: a signal voltage component d V,./dlproportional tome time rate of change of the aircraft velocity normal tothe desired flight path; and, voltage component dO/dt COS O Yproportional to the rate of change of aircraft pitch attitude rate, andwhich is also insensitive to the time rate of change of the aircraftsvelocity tangential to the flight path dV /dz. An additional versinesignal 136 derived in a manner well-known in the state of the art isgenerated from roll angle sensor 140 to compensate the accelerometersensor 139 and eliminate the effects of the versine term g (l-COS ICOS4)) inherent in a body mounted accelerometer. The output signal voltagefrom accelerometer circuit 1 is processed through lag filter 17 whichprovides an output signal voltage which is proportional to time rate ofchange of aircraft pitch attitude dB/a'! and velocity normal to flightpath, V A second circuit path is provided in series circuit betweenaccelerometer 1 and system output terminal 15 by means of lag filter 18connected in parallel with lag filter 17 to provide a further outputsignal voltage which is proportional to the output of accelerometer 1normal to flight path. In this manner, the critical damping termsnecessary for stability of the aircraft when flying an approach andlanding are derived from the single source (accelerometer 1) therebyincreasing the reliability of the system by reducing the number ofcritical component sensors necessary in the achievement of safestability margins.

The pitch rate voltage signal source utilized is pitch rate detector 3which is coupled in series circuit through band pass filters 1 11 andsumming resistor 1 13 to summingjunction 10 to provide an additionaldamping term in the system output signal voltage at output terminal 15by summation through summing junction 10. This damping term is notcritical in affecting aircraft or flight path stability.

A further feature of the presently described pitch axis control systemof FIG. 10 provides a unique and novel means of allowing glide slopeacquisition maneuvers at substantially any distance from the landingrunway (or from substantially any altitude above the runway). Thisaspect of the system hereinafter described has important considerationsin connection with noise abatement approaches wherein it is desirable toacquire the glide slope zero plane as close to the desired landing pointon the runway as possible to avoid long approaches over populated areas.ln this respect, the system embodiment of FIG. 10 utilizes a verticalbeam sensing means 66 (e.g., a threshold detector) which is locateddownstream in terms of signal processing from the gain programmercircuit 11. The gain programmer circuit 11 varies the glide slope errorsignal gain path (which includes the coupling of glide slope errordetector 4, gain programmer circuit 11, lag filter means (comprising alow pass filter) and summing resistor 117 in series circuit path tosumming junction 119) to convert the angular glide slope error signalvoltage from glide slope detector 4 into an error signal voltage whichis proportional to distance from the glide slope zero plane. In thismanner, the response of the system of FIG. 10 to glide slope errors ismaintained constant at substantially any altitude down to that altitudelevel at which the programmer output signal voltage from programmercircuit 11 is programmed to zero immediately prior to flaring of theaircraft. Since vertical beam sensor 66 is coupled in circuit betweenthe glide slope gain programmer circuit 11 and the system outputterminal 15 (downstream of the glide slope gain programmer circuit 11 interms of signal processing) the present system of FIG. is can maintain asubstantially constant distance from the glide slope zero plane forsystem activation irrespective of the distance from the runway that thesystem is engaged. This means that for low altitude glide slopeacquisitions, the vertical beam sensor 66 detection threshold isexceeded for greater error output signal voltages from glide slope errordetector 4 than it does for higher altitude glide slope acquisitionswhich results in an aircraft maneuver and flight path performance whichis substantially identical for both high and low altitude captures. Thisunique and novel feature allows the present automatic approach andlanding pitch axis control system to be utilized in the above manner fornoise abatement approaches heretofore not possi- A unique and novelflare command is provided by the present system which is switchless andprovides tigher control of landing dispersions along the runway. Theflare command signal voltage at circuit connection 19 in the seriescircuit comprising: altitude above terrain detector 5 connecting throughlimiter circuit 6 (comprising voltage limiting means, e.g., a saturatedamplifier), the parallel combination of altitude rate circuit 21comprising a high pass filter and altitude path displacement circuit 20comprising a summing resistor V proportional to displacement gain tosumming junction 7, asymmetrical limiter circuit 125, circuit connection19, summing resistor 127 to adder 129, adder 119, summing junction 10,and amplifier means to system output terminal 15, is a flare commandhaving a flare point and a touchdown rate of descent command which arevaried automatically by mechanization of the control laws to providetight control of the aircraft landing dispersions due to environmentalconditions such as winds, terrain and varying aircraft flight parametersuch as gross weight, center of gravity, flare configuration, and speed.The output voltage of the altitude above terrain detector 5 is limitedby limiter circuit 6 at an altitude so that large irregularities interrain distant from the normal flare point of the aircraft approachingthe landing runway do not affect the flare computation. Above thepredetermined altitude for which the limiter is set, the output oflimiter circuit 6 is a fixed parameter, i.e., not varying with time. Theoutput voltage from altitude rate circuit 21 is zero since the inputvoltage to circuit 21 is not time varying, and

the input voltages to summing junction 7 comprise the predeterminedlevel output voltage from voltage limiter circuit means 6 coupledthrough altitude path displacement circuit 20 which providesdisplacement path voltage amplification and the zero voltage output ofaltitude rate circuit 21. The input of the flare command signal voltageat lead 22 transmitted to summing junction is limited by asymmetricallimiter circuit 125 such that for positive summation of the inputvoltages to summing junction 7, no change in the output level of limitercircuit 125 can occur.

As the aircraft descends below the altitude at which the input voltageto limiter circuit 6 causes saturation the output voltage of limitercircuit 6 decreases in a manner proportional to altitude above theterrain. The output voltage from rate circuit 21 senses the rate ofchange of altitude with time and when the sum of the output voltages ofrate circuit 21 and altitude displacement circuit coupled into summingjunction 7 is negative in polarity, the output voltage 19 fromasymmetrical limiter circuit 125 is a command signal voltage 14 on lead22 coupled to summing junction 10 representative of a decreasingaltitude rate command. The preceding circuit feature enables an aircraftwhich is descending at a high sink rate to begin to command a flaremaneuver sooner than an aircraft descending at a lower voltage begins todecrease lf the aircraft begins to' float (i.e., approaches zero sinkrate) at some altitude above the runway, the summed output voltage fromsumming junction 7 decreases due to the decreasing voltage from rateoutput circuit 21 hence decreasing flare command called for by the flarecommand signal voltage on lead 22 thereby causing the aircraft toincrease its rate of sink for reducing touchdown dispersion.

The present system control laws discussed earlier as implemented in thepresent system embodiment allow generation of a switchless flare commandnot susceptible to switch failure prohibiting flare and further allowtouchdown dispersions due to environmental and air craft parameters tobe minimized in the manner hereinbefore discussed.

In addition to the preceding, the present system includes circuitfeatures which generate an automatic goaround command signal voltage atthe output of adder 23 which is generated in a manner such that as theaircraft enters the flare region an additional go-around command signalvoltage is generated as a portion of the switchless flare command signalvoltage present on lead 22 to reduce the altidude loss during thegoaround maneuver. The features of the go-around circuitry which areunique in the present system are that all of the same circuit componentsutilized in conducting the approach are utilized which are already knownoperative prior to initiation of go-around. Activation of the go-around(G/A) switch 24 by the pilot causes switch 14 to move to the openposition (shown by dotted line) thereby reducing the output of the glideslope gain programmer circuit 11 to zero and further causing switch 13to close (shown by the dotted line). Closing switch 13 connects togethersumming junctions 8 and 23 which results in conversion of glide slopeintegrator circuit 9 into a lag circuit through gain 142 with timeconstant -r'l/K Since the output voltage from the glide slope integratorcircuit 9 is proportional to the actual descent rate of the aircraft andis representative of a fly down command, the closing of switch 23 causessubsequent washout" elimination in the output of integrator circuit to aresultant zero fly down command thereby causing an error signal to begenerated at summing junction 10 to cause the aircraft to break its sinkrate and assume level flight. If the aircraft is in the flare region,the flare command signal voltage present on lead 22 will command theaircraft to climb to an altitude equivalent to the previously referredto flare initiation altitude and then maintain that altitude. Theseunique circuit features provide a go-around maneuver which is fail safein the sense that it uses known to be operating components and does notrequire the intro duction of a second signal source to initiate thegoaround maneuver (as is the case in the FIG. 1 system representative ofthe prior art), place the aircraft in level flight and maintain analtitude above the flare region.

A predetermined voltage level is provided by goaround bias source 26which is summedinto summing junction 23 thereby causing the outputsignal voltage of glide slope integrator circuit 9 to command a climbrate when switch 13 is closed.

Those skilled in the art will appreciate the important and significantfeature of the present system which provides a go-around command whichis fail safe in that it cannot inhibit a normal flare of the aircraft ifit fails and the further important feature that the system cannot causea nose down hardover situation as a result of failed components withinthe go-around circuitry and the further feature that the initialgo-around maneuver utilizes the same signal generating means alreadyknown to be operative.

FIG. 12 is a plot of actual airplane performance during a glide slopeacquisition, track and subsequent pilot initiated automatic go-aroundfor a Boeing 747. As can be seen the principles of this inventionprovide smooth and accurate acquisition of the glide slope zero plane(represented by zero glide slope error on the plot) and accuratetracking of the glide slope zero plane, and smooth and accuratego-around maneuver.

FIG. 12 includes automatic landing. Again the glide slope trackingaccuracy prior to flare can be measured to be inches by those skilled inthe art. The automatic nose lowering feature after touchdown can be seento be smooth providing the pilot the opportunity to keep his attentionon the task of stopping the airplane while the autopilot performs thetask of keeping the airplane on the ground.

What is claimed is:

1. In a pitch axis control system having an output terminal, flarecommand signal voltage generating means comprising:

altitude above terrain detector means;

a limiter circuit, a parallel connected altitude rate circuit andaltitude path' displacement circuit, asymmetrical limiter circuit means,and amplifier circuit means connected in series circuit between saidaltitude above terrain detector means and said output terminal therebyproviding a variable tangential flare point and a variable touchdownbias command as a function of time.

2. The invention according to claim 1 wherein a nose lowering commandvoltage is generated subsequent to touchdown for cancelling saidtouchdown bias command.

3. The invention according to claim 1 further comprising means forgenerating an automatic go-around command signal voltage in addition tosaid flare command signal voltage when the aircraft enters the flareregion to reduce the altitude loss during the go-around maneuver.

4. A pitch axis control system for providing a flight path commandsignal at the system output terminal comprising in combination:

glide slope error detector means;

gain programmer circuit means;

first adder circuit means; amplifier circuit means;

said gain programmer circuit means and said first adder circuit meanscoupled between said glide slope error detector means and the inputterminal of said amplifier means, the output terminal of said amplifiermeans coupled to said system output terminal;

surface feedback measuring means;

first means coupling said surface feedback means between the controlsurfaces and said system output terminal;

glide slope integrator circuit means coupled between said first addercircuit means and said output terminal;

altitude above terrain detector means;

flare command signal generating means including asymmetrical limitercircuit means coupled between said altitude above terrain detector meansand said first adder circuit means, and go-around switching means forcoupling a go-around bias source to said glide slope integrator circuit.

5. The invention according to claim 4 further comprising accelerationnormal to flight path detector means and altitude rate detector meanscoupled to said VIE'RUPEIIPIIWV f ?ld amrvl le means- 6. The inventionaccording to claim 5 further comprising pitch rate detector means andband pass filter means coupled in series circuit to said input terminalof said amplifier. if I g 7. A pitch axis control system for providing acontrol surface error command signal at a system output terminal toposition the control surfaces of an aircraft comprising in combination:

glide slope detector means, gain programmer circuit means, first lagcircuit means, first adder circuit means, and amplifier circuit means;said gain programmer circuit means, said first lag circuit means, andsaid first adder circuit means coupled in series circuit between saidglide slope error detector means and the input of said amplifier circuitmeans, the output of said amplifier circuit means coupled to said systemoutput terminal;

control surface feedback measuring means coupled between said controlsurfaces and said first adder circuit means;

acceleration normal to flight path detector means,

second adder ir cuit me ans; said accelera ion normal to flight pathdetector means coupled to said second adder circuit means;

roll attitude detector means,

versine generator means,

said versine generator means coupled in series circuit between said rollattitude detector means and said second adder circuit means;

first wash-out circuit means,

third adder circuit means,

said first wash-out circuit means coupled between said second addercircuit means and said third adder circuit means;

altitude rate detector means coupled to said third adder circuit means;

second lag circuit means,

said second lag circuit means coupled between said third adder circuitmeans and said first adder means;

altitude above terrain detector means,

first limiter circuit means,

fourth adder circuit means,

said limiter circuit means coupled between said altitude above theground detector means and said fourth adder circuit means;

first series switching means coupled between said altitude above terraindetector means and said gain programmer circuit means;

second wash-out circuit means coupled between said first limiter circuitmeans and said fourth adder cir cuit means;

second limiter circuit means coupled between said fourth adder circuitmeans and said first adder circuit means;

second series switching means,

fifth adder circuit means, and

integrator circuit means.

said fifth adder circuit means and said integrator circuit means coupledin series circuit between said second series switching means and saidfirst adder circuit means, said second series switching means connectedto said system output terminal prior to system engage, and to said gainprogrammer circuit means during system engage;

gain circuit means,

sixth adder circuit means, and

third series switching means,

said gain circuit means and said sixth adder circuit means coupled inseries circuit between the output of said integrator circuit means andsaid third series switching means;

go-around bias circuit means coupled to said sixth adder circuit means;

go-around switch means,

said go-around switch means when activated causing said first switchingmeans to reduce the output of said gain programmer circuit means to zeroand further causing said third series switching means to connect theoutput of said sixth adder circuit means to said fifth adder circuitmeans.

8. The invention according to claim 7 further com prising further lagcircuit means coupled between said first wash-out circuit means and saidfirst adder circuit means. 9. The invention according to claim 8 furthercomprising pitch rate detector means and band pass circuit means coupledin series circuit to said first adder circuit means.

10. In combination in a pitch axis control system for producing a signalrepresentative of the flight path error of an aircraft at an outputterminal during a glide slope approach and flare-out to a landing orduring a go-around manuever:

first means for generating a first signal representative of the altituderate of said aircraft, said first signal referenced to the aircraftflight path or short period maneuvering and to the aircraft verticalrate of descent or ascent for long term manuevering;

second means for selectively generating a second signal at the output ofan integrator circuit representative of commanded vertical rate ofdescent during the approach and flare-out phase of a landing or ofcommanded vertical rate of ascent during the go-around phase of anapproach;

third means for producing a third signal above a first predeterminedaltitude representative of deviation from glide slope zero plane duringthe approach phase of a landing;

fourth means for generating a fourth signal below a second predeterminedaltitude representative of the desired change in said commanded verticalrate of descent during the flare-out phase of a landing;

fifth means for combining said third signal and said fourth signal withsaid first signal and said second signal to produce said flight patherror signal at the output terminal.

11. The invention according to claim 10 further comprising sixth meansincluding longitudinal actuator,

1. In a pitch axis control system having an output terminal, flarecommand signal voltage generating means comprising: altitude aboveterrain detector means; a limiter circuit, a parallel connected altituderate circuit and altitude path displacement circuit, asymmetricallimiter circuit means, and amplifier circuit means connected in seriescircuit between said altitude above terrain detector means and saidoutput terminal thereby providing a variable tangential flare point anda variable touchdown bias command as a function of time.
 2. Theinvention according to claim 1 wherein a nose lowering command voltageis generated subsequent to touchdown for cancelling said touchdown biascommand.
 3. The invention according to claim 1 further comprising meansfor generating an automatic go-around command signal voltage in additionto said flare command signal voltage when the aircraft enters the flareregion to reduce the altitude loss during the go-around maneuver.
 4. Apitch axis control system for providing a flight path command signal atthe system output terminal comprising in combination: glide slope errordetector means; gain programmer circuit means; first adder circuitmeans; amplifier circuit means; said gain programmer circuit means andsaid first adder circuit means coupled between said glide slope errordetector means and the input terminal of said amplifier means, theoutput terminal of said amplifier means coupled to said system outputterminal; surface feedback measuring means; first means coupling saidsurface feedback means between the control surfaces and said systemoutput terminal; glide slope integrator circuit means coupled betweensaid first adder circuit means and said output terminal; altitude aboveterrain detector means; flare command signal generating means includingasymmetrical limiter circuit means coupled between said altitude aboveterrain detector means and said first adder circuit means, and go-aroundswitching means for coupling a go-around bias source to said glide slopeintegrator circuit.
 5. The invention according to claim 4 furthercomprising acceleration normal to flight path detector means andaltitude rate detector means coupled to said input terminal of saidamplifier means.
 6. The invention according to claim 5 furthercomprising pitch rate detector means and band pass filter means coupledin series circuit to said input terminal of said amplifier.
 7. A pitchaxis control system for providing a control surface error command signalat a system output terminal to position the control surfaces of anaircraft comprising in combination: glide slope detector means, gainprogrammer circuit means, first lag circuit means, first adder circuitmeans, and amplifier circuit means; said gain programmer circuit means,said first lag circuit means, and said first adder circuit means coupledin series circuit between said glide slope error detector means and theinput of said amplifier circuit means, the output of said amplifiercircuit means coupled to said system output terminal; control surfacefeedback measuring means coupled between said control surfaces and saidfirst adder circuit means; acceleration normal to flight path detectormeans, second adder circuit means; said acceleration normal to flightpath detector means coupled to said second adder circuit means; rollattitude detector means, versine generator means, said versine generatormeans coupled in series circuit between said roll attitude detectormeans and said second adder circuit means; first wash-out circuit means,third adder circuit means, said first wash-out circuit means coupledbetween said second adder circuit means and said third adder circuitmeans; altitude rate detector means coupled to said third adder circuitmeans; second lag circuit means, said second lag circuit means coupledbetween said third adder circuit means and said first adder means;altitude above terrain detector means, first limiter circuit means,fourth adder circuit means, said limiter circuit means coupled betweensaid altitude above the ground detector means and said fourth addercircuit means; first series switching means coupled between saidaltitude above terrain detector means and said gain programmer circuitmeans; second wash-out circuit means coupled between said first limitercircuit means and said fourth adder circuit means; second limitercircuit means coupled between said fourth adder circuit means and saidfirst adder circuit means; second series switching means, fifth addercircuit means, and integrator circuit means. said fifth adder circuitmeans and said integrator circuit means coupled in series circuitbetween said second series switching means and said first adder circuitmeans, said second series switching means connected to said systemoutput terminal prior to system engage, and to said gain programmercircuit means during system engage; gain circuit means, sixth addercircuit means, and third series switching means, said gain circuit meansand said sixth adder circuit means coupled in series circuit between theoutput of said integrator circuit means and said third series switchingmeans; go-around bias circuit means coupled to said sixth adder circuitmeans; go-around switch means, said go-around switch means whenactivated causing said first switching means to reduce the output ofsaid gain programmer circuit means to zero and further causing saidthird series switching means to connect the output of said sixth addercircuit means to said fifth adder circuit means.
 8. The inventionaccording to claim 7 further comprising further lag circuit meanscoupled between said first wash-out circuit means and said first addercircuit means.
 9. The invention according to claim 8 further comprisingpitch rate detector means and band pass circuit means coupled in seriescircuit to said first adder circuit means.
 10. In combination in a pitchaxis control system for producing a signal representative of the flightpath error of an aircraft at an output terminal during a glide slopeapproach and flare-out to a landing or during a go-around manuever:first means for generating a first signal representative of the altituderate of said aircraft, said first signal referenced to tHe aircraftflight path or short period maneuvering and to the aircraft verticalrate of descent or ascent for long term manuevering; second means forselectively generating a second signal at the output of an integratorcircuit representative of commanded vertical rate of descent during theapproach and flare-out phase of a landing or of commanded vertical rateof ascent during the go-around phase of an approach; third means forproducing a third signal above a first predetermined altituderepresentative of deviation from glide slope zero plane during theapproach phase of a landing; fourth means for generating a fourth signalbelow a second predetermined altitude representative of the desiredchange in said commanded vertical rate of descent during the flare-outphase of a landing; fifth means for combining said third signal and saidfourth signal with said first signal and said second signal to producesaid flight path error signal at the output terminal.
 11. The inventionaccording to claim 10 further comprising sixth means includinglongitudinal actuator means coupled between said fifth means and thecontrol surfaces of said aircraft.